The statements in this section merely provide background information related to the present disclosure and may not constitute prior art.
More particularly, it relates to a nacelle including at least one nacelle element having a wall delimiting a main flow channel, said channel having a variable flow passage section with at least one first area of the channel, the passage section of which is larger than that of at least one second area of the channel.
An airplane is driven by several turbojet engines each housed in a nacelle also harboring a set of ancillary actuation devices relating to its operation and ensuring various functions when the turbojet engine is operating or at a standstill. These ancillary actuation devices notably comprise a mechanical system for actuating thrust reversers.
With reference to FIG. 1, a nacelle N generally has a tubular structure comprising an air intake section SE1 upstream from the turbojet engine (not illustrated), a middle section SE2 intended to surround a fan S of the turbojet engine, a downstream section SE3 generally harboring thrust reversal means and intended to surround the combustion chamber of the turbojet engine, and is generally completed by an outlet section SES4, the outlet of which is located downstream from the turbojet engine.
The air intake section SE1 appears as an inlet structure element 1 of the nacelle N, located upstream from the turbojet engine and from the fan S of the turbojet engine, and this inlet structure element 11 has an internal tubular wall 10 delimiting a channel 11 for inflow of air into the nacelle N.
The middle section SE2 appears as a middle structure element 2, located upstream from the turbojet engine and surrounding the fan S, and this middle structure element 2 has an internal tubular wall 20 delimiting, with the fan S, a ring shaped air circulation channel 21.
The downstream section SE3 includes two concentric downstream structure elements 30, 31 located downstream from the fan S and surrounding the turbojet engine, i.e., a downstream external structure element 30, a so called Outer Fixed Structure (OFS), and a downstream internal structure element 31, a so called Inner Fixed Structure (IFS).
This downstream external structure element 30 has an internal tubular wall 32 and this downstream internal structure element 31 has an external ring shaped wall 33, wherein these tubular walls 32 and 33 delimit between them a ring shaped flow channel 34, also called a vein, intended for channelling the cold air flow circulating on the outside of the turbojet engine.
When operating, the turbojet engine of the dual flow type generates via the rotating blades of the fan S, a hot air flow (also called primary flow) from the combustion chamber of the turbojet engine, and a cold air flow (secondary flow) which circulates on the outside of the turbojet engine in the ring shaped flow channel 34.
The outlet section SE4 includes at least one gas ejection nozzle element 4, located downstream from the turbojet engine, wherein this nozzle element 4 includes at least one ring shaped wall 40 delimiting an air outflow channel 41 out of a nacelle N. In this way, the primary and secondary flows are ejected from the engine from the rear of the nacelle N, via said or these outlet channels 41.
In a nacelle said to be with separate flows, the primary and secondary flows are not mixed, i.e. they are separated at the outlet and circulate in primary and secondary nozzle elements respectively. In a so called nacelle with mixed flows, the primary and secondary flows are mixed downstream from the engine within a common nozzle element.
In order to limit sound pollution of airplanes and as illustrated in FIG. 1, it is conventional to at least partly cover the walls of the different nacelle elements with acoustic damping panels 5.
FIG. 2 illustrates an exemplary acoustic damping panel 5, of the sandwich composite panel type, including:                a structuring skin 50, of the so called “solid” air-proof skin type, attached against the wall of the relevant nacelle element;        a perforated skin 51, of the so called “acoustic” air-proof skin type at least partly delimiting the flow channel in this nacelle element;        a core 52 inserted between the structuring skin 50 and the perforated skin 51, said core 52 being made in an acoustic absorption material of the type having numerous internal cavities.        
Such panels form acoustic resonators capable of “trapping” the noise and therefore attenuating sound emissions towards the outside of the nacelle.
In the embodiment illustrated in FIG. 2, the core 52 is made in a cellular structure material of the honeycomb type. This material includes a plurality of honeycomb cells delimited by side walls, each honeycomb cell communicating with its neighbors. Such panels with a honeycomb structure are already known, notably from international applications WO 2008/113904 and WO 2009/066036, and will not be described in further details.
In an embodiment not shown, notably described in French patent applications FR 2 938 014 and FR 2 934 641, the core is made in a porous material; by “porous material” is meant an open material (i.e. having many communicating cavities) appearing as a foam or in an expanded form, or as a felt, or as an aggregate of small-size elements such as beads.
According to the architecture of the nacelle, certain of the aforementioned channels have a variable flow passage section with at least one first area of the channel, the passage section of which is greater than that of at least one second area of the channel. Given that the distribution of the pressure in a channel depends on the flow passage section, these channels have high pressure areas (associated with large passage sections) and low pressure areas (associated with small passage sections).
FIG. 3 illustrates the change in the flow passage section A (in m2) in the inlet channel 11 versus the longitudinal position X (in cm) along the middle axis AA′ of the nacelle N. At least two distinct areas Z1, Z2 are distinguished, i.e.:
a first small section area Z1 close to the air intake neck C, corresponding to an operating low pressure and high speed area on the airplane; and
a second large section area Z2, close to the fan, corresponding to an operating high pressure and low speed area on the airplane.
FIG. 4 illustrates on the curves C1 to C3, several examples of the change in the section A, based on the section Acol taken at the neck C, for letting through the flow into a ring shaped channel versus the longitudinal position X (in m) along the middle axis AA′ of the nacelle N. The curves Rint and Rext illustrate the change in the internal and external radii of the relevant ring shaped channel versus the longitudinal position X. At least two distinct areas Z1, Z2 are distinguished, i.e.:
a first small section area Z1, corresponding to an operating low pressure and high speed area on the airplane; and
a second large section area Z2 corresponding to an operating high pressure and low speed area on the airplane.
This ring shaped channel may be located in the downstream section SE3 or the outlet section SE4 of a nacelle N. For example, this ring shaped channel may correspond to the ring shaped channel 34, or vein, located between the downstream external structure element 30 (Outer Fixed Structure) and the downstream internal structure element 31 (Inner Fixed Structure). This ring shaped channel may also be located in a nozzle element, such as for example in a primary or secondary nozzle element for a nacelle with separate flows, or in a common nozzle element for a nacelle with mixed flows.
However, on the walls delimiting the different channels of the nacelle, the turbulent limiting layers of the flow generate pressure losses by friction. The impact of these pressure losses on the performance of the turbojet engine and therefore on the fuel consumption is significant. The applicants also noticed a loss of about 1.5% of the fuel consumption because of these frictional processes in the secondary nozzle element of a nacelle with separate flows.
The state of the art may also be illustrated by the teachings of patent applications US 2009/140104 A1, EP 1 517 022 A2 and U.S. Pat. No. 3,572,960 A.
The patent application US 2009/140104 A1 discloses a nacelle including blowing means intended to inject a tangential airflow into the internal volume of the nacelle, wherein these blowing means include at least one air inflow conduit capable of sampling the tangential air flow at the compressors of the turbojet engine, so as to guide the air flow in the nacelle, so as to be maintained parallel to the longitudinal axis of the nacelle and to suppress possible separations of air streams in the nacelle.
Patent application EP 1 517 022 A2 discloses a method for attenuating the turbojet engine noise consisting of increasing the flow velocity locally, reducing the limiting layer and the associated turbulences and optimizing refraction and absorption of sounds with an acoustic cover, wherein the increase in this velocity may be provided by means of a bypass conduit which sucks up and blows air into the nacelle by a difference in pressure.
Patent application U.S. Pat. No. 3,572,960 A discloses the placement of one or several conduits inside a fan cowl, between outlet rectifying fins and inlet directing fins positioned in the main flow channel, in order to at least substantially remove the wakes at the trailing edges of the inlet directing fins, and therefore reduce the noise.